Ceramic matrix composite airfoil repair

ABSTRACT

Methods for repairing composite components are provided. For instance, one exemplary method includes machining an interlocking feature into an existing component. A replacement material or core having an interlocking feature complementary to the interlocking feature of the existing component is then joined with the component. The interlocking features interlock to form a joint. The joint is then overlaid with one or more plies to rebuild the outer surface of the component and seal the joint. A bonding process can be used to chemically bond the newly joined parts together. Repaired composite components are also provided.

FIELD

The present subject matter relates generally to repairing components ofgas turbine engines. More particularly, the present subject matterrelates to repair of composite airfoils of gas turbine engines.

BACKGROUND

A gas turbine engine generally includes a fan and a core arranged inflow communication with one another. Additionally, the core of the gasturbine engine generally includes, in serial flow order, a compressorsection, a combustion section, a turbine section, and an exhaustsection. In operation, air is provided from the fan to an inlet of thecompressor section where one or more axial compressors progressivelycompress the air until it reaches the combustion section. Fuel is mixedwith the compressed air and burned within the combustion section toprovide combustion gases. The combustion gases are routed from thecombustion section to the turbine section. The flow of combustion gasesthrough the turbine section drives the turbine section and is thenrouted through the exhaust section, e.g., to atmosphere.

The turbine section includes one or more stages of a plurality ofstationary nozzle airfoils and a plurality of blade airfoils attached toa rotor that is driven by the flow of combustion gases against the bladeairfoils. The turbine section may have other configurations as well.Likewise, one or more of the compressors of the compressor section caninclude a plurality of stationary nozzle airfoils and a plurality ofblade airfoils attached to a rotor that is driven by the turbinesection. In addition, other sections of the gas turbine engine caninclude airfoils as well, such as e.g., the fan. Such airfoils typicallyhave complex geometries. For instance, airfoils can have relatively thinsections at their trailing edges and tips. Moreover, such airfoilshaving complex geometries can be formed from a composite laminate (i.e.,one or more layers of composite material). For instance, compressornozzles and blades can be formed from a polymer matrix composite (“PMC”)material and the turbine nozzles and blades can be formed from a ceramicmatrix composite (“CMC”) material.

Composite airfoils deteriorate over their service lives, and thus insome instances, require repair. The tips of blades as well as thetrailing edges of nozzles typically wear the fastest. Repairing theseareas has been challenging in the past due to the relatively thin crosssections of these areas. Conventional methods for repairing airfoilshave included attaching new plies to the damaged area. However, suchrepair methods have often led to distorted parts, as the thin sectionsof the airfoil offer little structure to which the new plies can attachand align with the existing structure. Other conventional methods haveincluded brazing a replacement material to the existing airfoil. Suchmethods typically require a melt alloy for bonding the parts together,which may, for example, affect the mechanical properties of the airfoil.Also, brazing a replacement material onto the existing airfoil has leftthe brazed joint particularly vulnerable to tensile and shear loadsexperienced by the airfoil during operation of the gas turbine engine.As a result, airfoils formed by such conventional processes are subjectto faster rates of wear and deterioration than original airfoils andthus require frequent further repairs. In short, conventional methodsfor repairing airfoils, especially at the portions of the airfoils withrelatively thin cross sections, and resulting repaired airfoils havebeen unsatisfactory.

Accordingly, improved methods for repairing composite components wouldbe desirable. In particular, improved methods for repairing compositeairfoils for gas turbine engines would be useful. Further, compositeairfoils repaired by such improved methods would be advantageous.

BRIEF DESCRIPTION

Aspects and advantages of the invention will be set forth in part in thefollowing description, or may be obvious from the description, or may belearned through practice of the invention.

In one exemplary embodiment of the present disclosure, a method forrepairing an airfoil formed of a composite material is provided. Themethod includes machining an interlocking feature into the airfoil. Themethod also includes joining a replacement core with the airfoil. Thereplacement core includes an interlocking feature complementary to theinterlocking feature of the airfoil. When the replacement core is joinedwith the airfoil, the interlocking features of the replacement core andthe airfoil interlock to form a joint. The method further includesoverlaying the joint with one or more plies.

In another exemplary embodiment of the present disclosure, a componentfor a gas turbine engine formed from a composite material is provided.The component includes an airfoil defining a chord length extendingbetween a leading edge and a trailing edge of the airfoil. The airfoilincludes an interlocking feature extending along the chord length. Thecomponent also includes a replacement core defining a chord lengthextending between a leading edge and a trailing edge of the replacementcore. The replacement core includes an interlocking featurecomplementary to the interlocking feature of the airfoil and extendingalong the chord length of the replacement core. The interlocking featureof the airfoil and the interlocking feature of the replacement materialare interlocked to form a joint. The component also includes one or moreplies overlaying the joint. At least one of the one or more plies arebonded to the airfoil and the replacement material.

In a further exemplary embodiment of the present disclosure, a methodfor repairing a component formed from a composite material is provided.The method includes joining a replacement material with the component.The component includes an interlocking feature and the replacementmaterial includes an interlocking feature complementary to theinterlocking feature of the component. When the replacement material isjoined with the component, the interlocking features of the replacementmaterial and the component interlock to form a joint. The method alsoincludes overlaying the joint with one or more plies. The method furtherincludes bonding the component with the replacement material and the oneor more plies with the component and the replacement material.

These and other features, aspects and advantages of the presentinvention will become better understood with reference to the followingdescription and appended claims. The accompanying drawings, which areincorporated in and constitute a part of this specification, illustrateembodiments of the invention and, together with the description, serveto explain the principles of the invention.

BRIEF DESCRIPTION OF THE DRAWINGS

A full and enabling disclosure of the present invention, including thebest mode thereof, directed to one of ordinary skill in the art, is setforth in the specification, which makes reference to the appendedfigures, in which:

FIG. 1 provides a schematic cross-section view of an exemplary gasturbine engine according to various embodiments of the present subjectmatter;

FIG. 2 provides a perspective view of an exemplary turbine blade of agas turbine engine according to an exemplary embodiment of the presentsubject matter;

FIG. 3 provides an exemplary airfoil having a damaged region accordingto an exemplary embodiment of the subject matter;

FIG. 4 provides the airfoil of FIG. 3 with the damaged region removedand having an interlocking feature formed into the airfoil as well as areplacement core having an interlocking feature according to anexemplary embodiment of the present subject matter;

FIG. 5 provides the airfoil joined with the replacement core accordingto an exemplary embodiment of the present subject matter;

FIG. 6 provides the airfoil of FIG. 3 depicting plies covering a jointaccording to an exemplary embodiment of the subject matter;

FIG. 7 provides a cross-sectional view of the turbine blade of FIG. 3taken along line 7-7 of FIG. 6 according to an exemplary embodiment ofthe present subject matter;

FIG. 8 provides an airfoil having an interlocking feature configured asa dovetail configuration as well as a replacement core having aninterlocking feature configured as a dovetail configurationcomplementary to the interlocking feature of the airfoil according to anexemplary embodiment of the present subject matter;

FIG. 9 provides an airfoil having an interlocking feature configured asa finger configuration as well as a replacement core having aninterlocking feature configured as a finger configuration complementaryto the interlocking feature of the airfoil according to an exemplaryembodiment of the present subject matter;

FIG. 10 provides an airfoil having an interlocking feature configured asa fir tree configuration as well as a replacement core having aninterlocking feature configured as a fir tree configurationcomplementary to the interlocking feature of the airfoil according to anexemplary embodiment of the present subject matter;

FIG. 11 provides a flow diagram of an exemplary method according to anexemplary embodiment of the present subject matter; and

FIG. 12 provides a flow diagram of another exemplary method according toan exemplary embodiment of the present subject matter.

Repeat use of reference characters in the present specification anddrawings is intended to represent the same or analogous features orelements of the present invention.

DETAILED DESCRIPTION

Reference will now be made in detail to present embodiments of theinvention, one or more examples of which are illustrated in theaccompanying drawings. The detailed description uses numerical andletter designations to refer to features in the drawings. Like orsimilar designations in the drawings and description have been used torefer to like or similar parts of the invention. As used herein, theterms “first,” “second,” and “third” may be used interchangeably todistinguish one component from another and are not intended to signifylocation or importance of the individual components. The terms“upstream” and “downstream” refer to the relative direction with respectto fluid flow in a fluid pathway. For example, “upstream” refers to thedirection from which the fluid flows and “downstream” refers to thedirection to which the fluid flows. As used herein, the term “about”means within ten percent of the stated value.

Aspects of the present disclosure are directed to methods for repairingcomposite components, such as e.g., CMC airfoils of a gas turbineengine. For instance, one exemplary aspect of the present disclosure isdirected to a method for repairing an airfoil for a gas turbine engine.The exemplary method includes machining an interlocking feature into theairfoil. Either prior to or at the same time as machining theinterlocking feature, a damages region of the airfoil can also beremoved. A replacement material or core having an interlocking featurecomplementary to the interlocking feature of the airfoil is then joinedwith the airfoil. The complementary interlocking features can be, forexample, complementary dovetails. The interlocking features of theairfoil and the replacement core interlock to form a joint. Theinterlocking features assist in aligning the core with the airfoil andcan counteract tensile and shear loads experienced by the airfoil duringoperation of the gas turbine engine. The joint is then overlaid with oneor more plies to rebuild the outer surface of the airfoil, seal thejoint, and to provide further structural integrity to the airfoil. Thereplacement core provides a structural component to which the plies canattach. A bonding process can be used to chemically bond the newlyjoined parts together. The airfoil can further undergo additionalmachining processes to remove any excess ply material such that theairfoil is formed to a predetermined or desired shape. Aspects of thepresent disclosure are also directed to repaired composite components.

Referring now to the drawings, wherein identical numerals indicate thesame elements throughout the figures, FIG. 1 is a schematiccross-sectional view of a gas turbine engine in accordance with anexemplary embodiment of the present disclosure. More particularly, forthe embodiment of FIG. 1, the gas turbine engine is a high-bypassturbofan jet engine 10, referred to herein as “turbofan engine 10.” Asshown in FIG. 1, the turbofan engine 10 defines an axial direction A(extending parallel to a longitudinal centerline 12 provided forreference) and a radial direction R. In general, the turbofan 10includes a fan section 14 and a core turbine engine 16 disposeddownstream from the fan section 14.

The exemplary core turbine engine 16 depicted generally includes asubstantially tubular outer casing 18 that defines an annular inlet 20.The outer casing 18 encases, in serial flow relationship, a compressorsection including a booster or low pressure (LP) compressor 22 and ahigh pressure (HP) compressor 24; a combustion section 26; a turbinesection including a high pressure (HP) turbine 28 and a low pressure(LP) turbine 30; and a jet exhaust nozzle section 32. A high pressure(HP) shaft or spool 34 drivingly connects the HP turbine 28 to the HPcompressor 24. A low pressure (LP) shaft or spool 36 drivingly connectsthe LP turbine 30 to the LP compressor 22. In other embodiments ofturbofan engine 10, additional spools may be provided such that engine10 may be described as a multi-spool engine.

For the depicted embodiment, fan section 14 includes a fan 38 having aplurality of fan blades 40 coupled to a disk 42 in a spaced apartmanner. As depicted, fan blades 40 extend outward from disk 42 generallyalong the radial direction R. The fan blades 40 and disk 42 are togetherrotatable about the longitudinal axis 12 by LP shaft 36. In someembodiments, a power gear box having a plurality of gears may beincluded for stepping down the rotational speed of the LP shaft 36 to amore efficient rotational fan speed.

Referring still to the exemplary embodiment of FIG. 1, disk 42 iscovered by rotatable front nacelle 48 aerodynamically contoured topromote an airflow through the plurality of fan blades 40. Additionally,the exemplary fan section 14 includes an annular fan casing or outernacelle 50 that circumferentially surrounds the fan 38 and/or at least aportion of the core turbine engine 16. It should be appreciated thatnacelle 50 may be configured to be supported relative to the coreturbine engine 16 by a plurality of circumferentially-spaced outletguide vanes 52. Moreover, a downstream section 54 of the nacelle 50 mayextend over an outer portion of the core turbine engine 16 so as todefine a bypass airflow passage 56 therebetween.

During operation of the turbofan engine 10, a volume of air 58 entersturbofan 10 through an associated inlet 60 of the nacelle 50 and/or fansection 14. As the volume of air 58 passes across fan blades 40, a firstportion of the air 58 as indicated by arrows 62 is directed or routedinto the bypass airflow passage 56 and a second portion of the air 58 asindicated by arrows 64 is directed or routed into the LP compressor 22.The ratio between the first portion of air 62 and the second portion ofair 64 is commonly known as a bypass ratio. The pressure of the secondportion of air 64 is then increased as it is routed through the highpressure (HP) compressor 24 and into the combustion section 26, where itis mixed with fuel and burned to provide combustion gases 66.

The combustion gases 66 are routed through the HP turbine 28 where aportion of thermal and/or kinetic energy from the combustion gases 66 isextracted via sequential stages of HP turbine stator vanes 68 that arecoupled to the outer casing 18 and HP turbine rotor blades 70 that arecoupled to the HP shaft or spool 34, thus causing the HP shaft or spool34 to rotate, thereby supporting operation of the HP compressor 24. Thecombustion gases 66 are then routed through the LP turbine 30 where asecond portion of thermal and kinetic energy is extracted from thecombustion gases 66 via sequential stages of LP turbine stator vanes 72that are coupled to the outer casing 18 and LP turbine rotor blades 74that are coupled to the LP shaft or spool 36, thus causing the LP shaftor spool 36 to rotate, thereby supporting operation of the LP compressor22 and/or rotation of the fan 38.

The combustion gases 66 are subsequently routed through the jet exhaustnozzle section 32 of the core turbine engine 16 to provide propulsivethrust. Simultaneously, the pressure of the first portion of air 62 issubstantially increased as the first portion of air 62 is routed throughthe bypass airflow passage 56 before it is exhausted from a fan nozzleexhaust section 76 of the turbofan 10, also providing propulsive thrust.The HP turbine 28, the LP turbine 30, and the jet exhaust nozzle section32 at least partially define a hot gas path 78 for routing thecombustion gases 66 through the core turbine engine 16.

It will be appreciated that, although described with respect to turbofan10 having core turbine engine 16, the present subject matter may beapplicable to other types of turbomachinery. For example, the presentsubject matter may be suitable for use with or in turboprops,turboshafts, turbojets, industrial and marine gas turbine engines,and/or auxiliary power units.

In some embodiments, components of turbofan engine 10, particularlycomponents within hot gas path 78, such as components of combustionsection 26, HP turbine 28, and/or LP turbine 30, may comprise a ceramicmatrix composite (CMC) material, which is a non-metallic material havinghigh temperature capability. Of course, other components of turbofanengine 10, such as components of HP compressor 24, may comprise a CMCmaterial. Exemplary CMC materials utilized for such components mayinclude silicon carbide (SiC), silicon, silica, or alumina matrixmaterials and combinations thereof. Ceramic fibers may be embeddedwithin the matrix, such as oxidation stable reinforcing fibers includingmonofilaments like sapphire and silicon carbide (e.g., Textron's SCS-6),as well as rovings and yarn including silicon carbide (e.g., NipponCarbon's NICALON®, Ube Industries' TYRANNO®, and Dow Corning'sSYLRAMIC®), alumina silicates (e.g., Nextel's 440 and 480), and choppedwhiskers and fibers (e.g., Nextel's 440 and SAFFIL®), and optionallyceramic particles (e.g., oxides of Si, Al, Zr, Y, and combinationsthereof) and inorganic fillers (e.g., pyrophyllite, wollastonite, mica,talc, kyanite, and montmorillonite). For example, in certainembodiments, bundles of the fibers, which may include a ceramicrefractory material coating, are formed as a reinforced tape, such as aunidirectional reinforced tape. A plurality of the tapes may be laid uptogether (e.g., as plies) to form a preform component. The bundles offibers may be impregnated with a slurry composition prior to forming thepreform or after formation of the preform. The preform may then undergothermal processing, such as a cure or burn-out to yield a high charresidue in the preform, and subsequent chemical processing, such asmelt-infiltration or chemical vapor infiltration with silicon, to arriveat a component formed of a CMC material having a desired chemicalcomposition. In other embodiments, the CMC material may be formed as,e.g., a carbon fiber cloth rather than as a tape.

As stated, components that include a CMC material may be used within thehot gas path 78, such as within the combustion and/or turbine sectionsof engine 10. As an example, the combustion section 26 may include acombustor formed from a CMC material and/or one or more stages of one ormore stages of the HP turbine 28 may be formed from a CMC material.However, CMC components may be used in other sections as well, such asthe compressor and/or fan sections. In some embodiments, other hightemperature materials and/or other composite materials may be used toform one or more components of engine 10.

FIG. 2 provides an exemplary composite component depicted as a turbineblade for a gas turbine engine, such as e.g., one of the HP turbinerotor blades 70 of the turbofan engine 10 of FIG. 1. Although thecomposite component is depicted as a turbine blade for use in a turbineblade assembly, in other exemplary embodiments, the composite componentcan be a blade configured for use in a fan blade assembly, a compressorblade assembly, or any other suitable application. Moreover, for thisembodiment, the turbine blade 70 is formed from a CMC material, such ase.g., silicon carbide fibers embedded in a silicon carbide matrix(SiC/SiC). Although blade 70 is depicted as being formed from a CMCmaterial, in other exemplary embodiments, various blades of a gasturbine engine can include other matrix materials, such as epoxymaterials (e.g., for fans), polymer composites (e.g., for compressors),or any other suitable matrix material.

As shown in FIG. 2, the turbine blade 70 includes an airfoil 100 againstwhich the flow of hot combustion gases 66 (FIG. 1) are directed. Theairfoil 100 includes a leading edge 102, a trailing edge 104, a pressureside wall 106, and a suction side wall 108 opposite the pressure sidewall 106. The pressure side wall 106 is connected to the suction sidewall 108 at leading edge 102 and trailing edge 104. The airfoil 100further includes a tip 110 and a root 112 connected by pressure sidewall 106, suction side wall 108, leading edge 102, and trailing edge104. The airfoil 100 has a span S extending between the root 112 and thetip 110 of the airfoil 100 along the radial direction R and a chordlength CL extending between the leading edge 102 and the trailing edge104. A mid-span reference line RL is defined midway between the root 112and the tip 110 along the span S of the airfoil 100. As further shown inFIG. 2, the airfoil 100 is connected to a shank 80. The shank 80includes a platform 82 and a dovetail 84. The turbine blade 70 can bemounted to a turbine disk (not shown) by engaging the dovetail 84 withina slot having a complementary geometry (not shown). The airfoil 100 isconnected at its root 112 to the platform 82.

During operation of the gas turbine engine 100, various airfoils, suchas e.g., the turbine blade 70 of FIG. 2, compressor blades, otherturbine blades, compressor stator vanes, turbine nozzles or statorvanes, fan blades, etc. are subjected to extreme pressures and/ortemperatures, causing deterioration of the airfoils over time. Moreover,the airfoils can further be degraded or worn by foreign object debris(FOD). For example, the fan blades 40 are particularly vulnerable to FODas the fan 134 is positioned at the inlet 156 of the gas turbine engine100 (See FIG. 1). In some instances, damaged or deteriorated airfoilsare removed from the engine and undergo a repair process.

FIGS. 3 through 7 provide an exemplary method for repairing a damagedairfoil, such as e.g., the turbine blade 70 of FIG. 2. In particular,FIG. 3 provides an exemplary airfoil 100 having a deteriorated ordamaged region 114. FIG. 4 provides the airfoil 100 with the damagedregion 114 removed and depicting the airfoil 100 having an interlockingfeature 116. FIG. 4 further depicts a replacement material or core 130having an interlocking feature 144 configured to interlock with theinterlocking feature 116 of the airfoil 100 to form a joint 160. FIG. 5provides the airfoil 100 joined with the replacement core 130. FIG. 6provides the airfoil 100 joined with the replacement core 130 anddepicts plies 170 overlaying the joint 160. FIG. 7 provides across-sectional view of the airfoil 100 taken along line 7-7 of FIG. 6according to an exemplary embodiment of the present subject matter.

As shown in FIG. 3, the airfoil 100 has experienced significantdeterioration, and more particularly, the airfoil 100 has experiencedsignificant abrasion or wear along the tip 110 of the airfoil 100.Airfoils for gas turbine engines can experience a wide variety of damagetypes or failures, including microstructural change, cracks, abrasion,deformation, and entire breakages. Such deterioration negatively affectsengine performance and efficiency. As noted above, in some instances, itis desirable to repair such deteriorated airfoils.

As shown in FIG. 4, interlocking feature 116 is shown machined into theairfoil 100, and by machining the interlocking feature 116 into theairfoil 100, the damaged region 114 (FIG. 3) is removed. Theinterlocking feature 116 can be cut or machined into the airfoil 100 andthe damaged region 114 can be removed by any suitable material removalprocess, such as e.g., a cutting process. Although the damaged region114 is described as being removed at the same time as the machining ofthe interlocking feature 116 for this embodiment, in some embodiments,the damaged region 114 can be removed prior to forming the interlockingfeature 116 into the airfoil 100. In such embodiments, for example, thedamaged region 114 can be removed by making a straight line cut acrossthe chord length CL of the airfoil 100. The interlocking feature 116 canthen be machined into the newly formed straight edge. This mayfacilitate alignment of the cutting relative to the airfoil such thatthe interlocking feature 116 can be machined high accuracy.

Referring still to FIG. 4, for this embodiment, the interlocking feature116 machined into the airfoil 100 extends along the chord length CL ofthe airfoil 100. Moreover, for this embodiment, the interlocking feature116 of the airfoil 100 is a series of dovetails 118 spaced apart alongthe chord length CL. Each dovetail 118 has a thickness extending alongthe circumferential direction C between the pressure side 106 and thesuction side 108 (FIG. 2) of the airfoil 100.

In some embodiments, prior to machining, the airfoil 100 defines a spanS extending between root 112 and tip 110 of the airfoil 100 and mid-spanreference line RL is defined midway between root 112 and tip 110 (e.g.,as shown in FIG. 2). For this embodiment, when machining theinterlocking feature 116 into the airfoil 100, the interlocking feature116 is machined into the airfoil 100 between about the reference line RLand root 112. By machining the interlocking feature 116 between aboutthe reference line RL and root 112, there is a higher probability thatthe damaged region 114 has been completely removed and thus a degree ofsafety is achieved. In some instances, structural cracks in the airfoil100 can extend from the tip 110 and propagate along the span S theairfoil 100. Accordingly, machining the interlocking feature 116 betweenabout the reference line RL and root 112 can better ensure that thedamaged region 114 has been completely removed. In addition, bymachining the interlocking feature 116 between about the reference lineRL and root 112, as will be explained more fully below, the core span CSor length of the replacement core 130 is increased and thus there ismore attachment area for the one or more plies 170 to attach to thereplacement core 130. Thus, greater structural rigidity can be achieved.As used in the context of this paragraph, the span S is indicative of anoriginal span S of the airfoil without damage (e.g., as shown in FIG.2).

In some embodiments, prior to machining, the airfoil 100 defines a spanS extending between root 112 and tip 110 of the airfoil 100 and mid-spanreference line RL is defined midway between root 112 and tip 110 (e.g.,as shown in FIG. 2). For this embodiment, when machining theinterlocking feature 116 into the airfoil 100, the interlocking feature116 is machined into the airfoil 100 a distance from the tip 110 that isat least about twenty percent (20%) of the span S. By machining theinterlocking feature 116 at least a distance of twenty percent (20%) ofthe span S, it is ensured that the replacement core 130 has a sufficientattachment area for the one or more plies 170 to attach to thereplacement core 130. In yet some further embodiments, when machiningthe interlocking feature 116 into the airfoil 100, the interlockingfeature 116 is machined into the airfoil 100 a distance from the tip 110that is at least about ten percent (10%) of the span S, at least aboutfifteen percent (15%) of the span S, at least about thirty percent (30%)of the span S, or at least about forty percent (40%) of the span S. Asused in the context of this paragraph, the span S is indicative of anoriginal span S of the airfoil without damage (e.g., as shown in FIG.2).

As further shown in FIG. 4, replacement core 130 is positioned forjoining the existing airfoil 100. Replacement core 130 includes aleading edge 132, a trailing edge 134, a pressure side wall 136, and asuction side wall 138 (FIG. 7) opposite the pressure side wall 136. Thepressure side wall 136 is connected to the suction side wall at leadingedge 132 and trailing edge 134. Replacement core 130 further includes atip end 140 and a connection end 142 connected by pressure side wall136, suction side wall 138, leading edge 132, and trailing edge 134. Thereplacement core 130 has a core span CS extending between the tip end140 and the connection end 142 and a chord length CL extending betweenthe leading edge 132 and the trailing edge 134 of the replacement core130. As shown in FIG. 4, the chord length CL of the replacement core 130can vary along the core span CS. The replacement core 130 can be, forexample, any suitable composite material, such as e.g., a CMC, PMC, orother composite material. The replacement core 130 can be formed of thesame material as the airfoil 100 to which it is configured to beattached. In this way, the replacement core 130 can be formed of acomplementary material to the airfoil 100. For instance, if the airfoil100 is a PMC material configured for use in the compressor section of agas turbine engine, the replacement core 130 can likewise be formed of aPMC material. Similarly, if the airfoil 100 is a CMC material configuredfor use in the turbine section of a gas turbine engine, the replacementcore 130 can be formed of a CMC material.

In addition, with reference still to FIG. 4, the replacement core 130includes an interlocking feature 144 that is complementary to theinterlocking feature 116 of the airfoil 100. Accordingly, for thisembodiment, the interlocking feature 144 of the replacement core 130 isalso a series of dovetails 146 spaced apart along the chord length CL ofthe replacement core 130. For this embodiment, the airfoil 100 includesfour dovetail portions or rounded tenons and three slots or mortises.The replacement core 130, on the other hand, includes three dovetailportions or rounded tenons and four slots or mortises, with the endslots or mortises being a half slot or mortise.

Although the interlocking features 116, 144 are depicted as havingdovetail configurations, the interlocking features 116, 144 of theairfoil 100 and replacement core 130 can be any suitable mechanicalfeature that allows the airfoil 100 to interlock with the replacementcore 130. For instance, as shown in FIGS. 8 through 10, furtherexemplary embodiments of interlocking features 116, 144 are provided.

FIG. 8 provides airfoil 100 having an interlocking feature 116configured as a dovetail configuration 118 as well as a replacement corehaving an interlocking feature 144 configured as a dovetailconfiguration 146 complementary to the interlocking feature 116 of theairfoil 100. However, in the embodiment of FIG. 8, the tenons areflaring tenons as opposed to rounded tenons as shown in FIG. 4. Theslots or mortises are likewise configured to receive the flaring tenons.

FIG. 9 provides airfoil 100 having interlocking feature 116 configuredas a series of comb or finger projections 150 as well as replacementcore 130 having interlocking feature 144 configured as a series offinger projections 152 complementary to the finger projections 150 ofthe airfoil 100.

FIG. 10 provides airfoil 100 having interlocking feature 116 configuredas fir trees 154 as well as replacement core 130 having interlockingfeature 144 configured as fir trees 156. As shown, the fir trees 154,156 and slots or mortises spacing the fir trees apart from one anotheralong the respective chord length CL of the airfoil 100 and replacementcore 130 are complementary to one another. It will be appreciated thatinterlocking features 116, 144 can take other suitable configurations.

Returning to FIGS. 3 through 7 and particularly to FIG. 4, as shown, thereplacement core 130 defines a plurality of cooling holes 148. For thisembodiment, the cooling holes 148 are positioned proximate the trailingedge 134 of the replacement core 130. In other embodiments, coolingholes 148 can be positioned in any suitable location for cooling thereplacement core 130. For instance, cooling holes 148 can be positionedalong the tip end 140 of the replacement core 130. As the trailing edgeof airfoils can become particularly hot during operation of the engine,cooling holes 148 can be provided to receive a cooling fluid and coolsuch trailing edges. Composite airfoils, particularly CMC compositeairfoils, can be brittle and thus machining cooling holes into arelatively thin cross section of such airfoils can be challenging, ascracks and other damage to airfoil can occur during formation of suchcooling holes. By providing preformed cooling holes 148 in thereplacement core 130, the repaired airfoil can receive the benefit ofhaving one or more cooling passages without being subjected to machininginto the brittle outer layers of the airfoil.

As shown in FIG. 5, by interlocking the interlocking feature 116 of theairfoil 100 and the interlocking feature 144 of the replacement core130, the airfoil 100 can be connected to or otherwise joined to thereplacement core 130 to form joint 160. The interlocking features 116,144 provide for secure attachment of the placement core 130 to theairfoil 100 and also assist with alignment of the replacement core 130with the airfoil 100. In particular, for rotating airfoils, theinterlocking features 116, 144 counteract tensile forces by preventingthe replacement core 130 from flying radially outward as the blades arerotated about during operation and also counteract shear forces bypreventing the replacement core 130 from sliding along thecircumferential direction C relative to the airfoil 100 as the airfoilsare rotated about. For stationary airfoils, the interlocking features116, 144 counteract shear forces by preventing mass flows through theengine from shearing or sliding the replacement core 130 relative to theexisting airfoil 100. Moreover, the interlocking features 116, 144 alsocounteract tensile forces as well.

For this embodiment, to interlock the replacement core 130 with theairfoil 100, the rounded tenons of the dovetail configuration 118 of theairfoil 100 are inserted or slid into the slots or mortises of thedovetail configuration 146 of the replacement core 130. Likewise, therounded tenons of the dovetail configuration 146 of the replacement core130 are inserted or slid into the slots or mortises of the dovetailconfiguration 118 of the airfoil 100. By interlocking replacement core130 to the airfoil 100, the tensile and shear loads experienced by thereplacement core 130 during operation of the gas turbine engine can betransferred to the existing airfoil 100. Moreover, as described furtherbelow, the replacement core 130 provides a structure to which plies 170can be attached.

In some embodiments, after joining the airfoil 100 with the replacementcore 130, the joined components can optionally undergo one or morebonding processes such that the existing airfoil 100 chemically bondswith the replacement core 130. For example, the joined components can beinserted into an autoclave for a predetermined time or until the joinedcomponents reach a predetermined temperature.

As shown in FIGS. 6 and 7, the airfoil 100 is joined with thereplacement core 130 at joint 160 via their respective interlockingfeatures 116, 144 in a manner described above. After being joined, thejoint 160 is covered or overlaid with composite plies 170 to rebuild andrecontour the complex outer surfaces of the airfoil 100. In addition,the plies 170 are overlaid over the joint 160 to provide shear strengthto the repaired airfoil and act to seal the joint 160.

For this embodiment, the laminate of plies 170 overlaying the joint 160include a first ply layer 172, a second ply layer 174, and a third plylayer 176. Although three ply layers are shown in FIGS. 6 and 7, thelaminate of plies 170 can include more than three ply layers or lessthan the three ply layers. As shown particularly in FIG. 7, at least oneof the plies 170 overlaying the joint 160, e.g., one of the plies of thefirst ply layer 172, is connected to both the replacement core 130 andthe airfoil 100. In some embodiments, as shown in FIG. 7, one ply 170can overlay the joint 160 and can connect to both the replacement core130 and the airfoil 100 along the pressure side 136 of the airfoil 100and one ply 170 can overlay the joint 160 and can connect to both thereplacement core 130 and the airfoil 100 along the suction side 138 ofthe airfoil 100. As shown further in FIGS. 6 and 7, after the first plylayer 170 is laid over the joint 160, for this embodiment, plies 170forming the second ply layer 174 and the third ply layer 176 are thenoverlaid on top of the first ply layer 172. The layers of plies 170 canbe laid up such that the predetermined geometry (i.e., the desiredshape) of the airfoil 100 can be formed.

First ply layer 172 can be made up of multiple plies 170 or in someembodiments can be made of a single ply 170 (i.e., a ply can be wrappedaround the tip end 140 of the replacement core and can extend along thepressure and suction sides 136, 138 of the airfoil 100. As shown in FIG.7, for this embodiment, the first ply layer 172 is made up of multiplyplies 170. One ply 170 of the first ply layer 172 extends along thepressure side 136 and overlays joint 160. The ply 170 is attached to orconnected with both the existing airfoil 100 and the replacement core130. Proximate the tip end 140 of the replacement core 130, the ply 170is joined or mates with a second ply 170 of the first ply layer 172. Thesecond ply 170 extends along the suction side 138 and overlays joint160. The second ply 170 is attached to or connected with both theexisting airfoil 100 and the replacement core 130 as shown. In this way,the joint 160 is sealed by the plies 170 of the first ply layer 172, andin addition to the mechanically interlocking, the replacement core 130is securely attached to the airfoil 100 by the plies 170, particularlyafter the repaired airfoil undergoes a suitable bonding process asdescribed more fully herein. Likewise, the second ply layer 174 and thethird ply layer 176 can be made up of multiple plies 170 as shown inFIG. 7 or in some embodiments can be made of a single ply 170. Theadditional ply layers 174, 176 can further seal the joint 160, provideadditional structural rigidity to counteract tensile and shear loads,and can form the repaired airfoil to a predetermined geometry or desiredshape. There need not be the same amount of ply layers on the pressureand suctions sides 136, 138 of the airfoil 100. Moreover, the plies 170can have different lengths as shown in FIG. 7.

In some embodiments, to ensure secure attachment of the plies 170 to theairfoil 100 and the replacement core 130, one or more of the plies 170can overlay the joint 160 in the following manner. As shown in FIG. 6,the joint 160 defines a joint width W. The joint width W defines acenter C. In some embodiments, at least one of the plies 70 extends fromthe center C of the joint width W in a first direction D1 a distanceequal to the joint width W and extends from the center C of the jointwidth W in a second direction D2 a distance equal to the joint width W,the first direction D1 opposite the second direction D2.

After overlaying the plies 170 over joint 160, in the event there isexcess material, the plies 170 and/or other parts of the repairedairfoil can be machined such that the repaired airfoil is shaped to apredetermined geometry or desired shape of the airfoil 100. Any suitablemachining process can be used to machine the excess material from theplies 170, such as e.g., any suitable cutting process.

In yet other embodiments, after joining the airfoil 100 with thereplacement core 130 and thereafter covering the joint 160 with plies170, the joined components and now attached plies can collectivelyundergo one or more bonding processes such that the existing airfoil 100chemically bonds with the replacement core 130 and the plies 170. Inparticular, in some embodiments, after the plurality of plies 170 arelaid up or overlaid over the joint 160, the preform, reshaped airfoil iscured to produce a single piece, unitary composite component, which isthen fired and subjected to densification, e.g., siliconmelt-infiltration, to form a final unitary composite structure.

For instance, the repaired airfoil preform can be processed in anautoclave to produce a green state unitary repaired airfoil. Then, thegreen state repaired airfoil can be placed in a furnace to burn outexcess binders or the like and then can be placed in a furnace with apiece or slab of silicon and fired to melt infiltrate the repairedairfoil with at least silicon. More particularly, for the repairedairfoil formed from CMC plies of prepreg tapes that are produced asdescribed above, heating (i.e., firing) the green state component in avacuum or inert atmosphere decomposes the binders, removes the solvents,and converts the precursor to the desired CMC material. Thedecomposition of the binders results in a porous CMC body; the body mayundergo densification, e.g., melt infiltration (MI), to fill theporosity. In one example, where the green state repaired airfoil isfired with silicon, the repaired airfoil can undergo siliconmelt-infiltration. However, densification can be performed using anyknown densification technique including, but not limited to, Silcomp,melt infiltration (MI), chemical vapor infiltration (CVI), polymerinfiltration and pyrolysis (PIP), and oxide/oxide processes, and withany suitable materials including but not limited to silicon. In oneembodiment, densification and firing may be conducted in a vacuumfurnace or an inert atmosphere having an established atmosphere attemperatures above 1200° C. to allow silicon or other appropriatematerial or combination of materials to melt-infiltrate into thecomponent. The densified CMC body hardens to a final unitary CMCrepaired airfoil.

In some embodiments, the final unitary structure may be finish machined,e.g., to bring the structure within tolerance, to shape the repairedairfoil to a predetermined geometry or desired shape, and/or add anenvironmental barrier coating (EBC) to the unitary repaired airfoil,e.g., to protect the repaired airfoil from the hot combustion gases 66(FIG. 1). It will be appreciated that other methods or processes offorming composite components, such as unitary composite repairedairfoil, can be used as well.

FIG. 11 provides a flow diagram of an exemplary method according to anexemplary embodiment of the present subject matter. In particular, FIG.11 provides a flow diagram for a method for repairing an airfoil formedof a composite material, such as e.g., a CMC material.

At (302), the method (300) includes machining an interlocking featureinto the airfoil. For instance, the interlocking feature can be theinterlocking feature 116 shown and described herein. For example, theinterlocking feature 116 can be dovetails 118 (rounded as shown in FIG.4 or flaring as shown in FIG. 8), finger projections 150 as shown inFIG. 9, fir trees 154 as shown in FIG. 10, a combination of theforegoing, etc.

In some implementations, prior to machining the interlocking featureinto the airfoil, the airfoil defines a deteriorated region. In suchimplementations, during machining, the method includes removing thedeteriorated region from the airfoil. This can be done simultaneously orin steps. For instance, as noted above, the airfoil can first be cutalong a straight edge to remove the deteriorated region and to provideeasier access for cutting the interlocking feature into the airfoil.

In some further implementations, prior to machining, the airfoil definesa span extending between a root and a tip of the airfoil. In suchimplementations, a reference line is defined midway between the root andthe tip. During machining the interlocking feature into the airfoil, theinterlocking feature is machined into the airfoil between about thereference line and the root.

In yet other implementations, the airfoil defines a span extendingbetween a root and a tip of the airfoil, the span indicative of anoriginal span of the airfoil without damage or deterioration. In suchimplementations, the interlocking feature is machined into the airfoil adistance from the tip that is at least about twenty percent (20%) of thespan.

At (304), the method (300) includes joining a replacement core with theairfoil, wherein the replacement core comprises an interlocking featurecomplementary to the interlocking feature of the airfoil, and whereinwhen the replacement core is joined with the airfoil, the interlockingfeatures of the replacement core and the airfoil interlock to form ajoint. For example, the replacement core can be the replacement core 130as depicted and described herein. The replacement core 130 can includeinterlocking feature 144 illustrated and described herein as well. Theinterlocking feature 144 of the replacement core 130 can becomplementary to the interlocking feature 116 of the airfoil. Forexample, the interlocking features 116, 144 of the airfoil 100 and thereplacement core 130 can both have dovetail configurations 118, 146, asshown in FIG. 4. The joint formed by the interlocking features can bejoint 160 as shown in FIG. 5.

In some implementations, the airfoil defines a chord length extendingbetween a leading edge and a trailing edge and wherein the replacementcore defines a chord length extending between a leading edge and atrailing edge of the replacement core, and wherein the interlockingfeatures of the replacement core and the airfoil extend substantiallyalong the chord length. In some embodiments, the interlocking featuresof the replacement core and the airfoil extend along their entirerespective chord lengths. In some implementations, the interlockingfeatures of the replacement core and the airfoil extend along theirentire respective camber lines.

At (306), the method (300) includes overlaying the joint with one ormore plies. For instance, one or more plies 170 as shown in FIGS. 6 and7 can overlay the joint 160 formed by the interlocking features 116, 144of the airfoil 100 and replacement core 130, respectively. By overlayingplies 170 over the joint 160, the airfoil can be rebuilt or recontouredto a predetermined geometry or desired shape, the plies 170 can provideimproved shear and tensile strength to the repaired airfoil and act toseal the joint 160. In some implementations, any excess material of theplies 170 can be finish machined from the repaired airfoil such that theairfoil is formed to a predetermined geometry, the predeterminedgeometry being indicative of a desired shape of the airfoil. Moreover,in some implementations, to ensure a secure connection between thereplacement core 130 and the airfoil 100, at least one of the one ormore plies 170 overlaying the joint 160 is connected to both thereplacement core 130 and the airfoil 100.

In yet other implementations, the joint defines a joint width, andwherein at least one of the one or more plies extends from a center ofthe joint width in a first direction a distance equal to the joint widthand extends from the center of the joint width in a second direction adistance equal to the joint width, the first direction opposite thesecond direction.

In some implementations, after overlaying the joint with one or moreplies, the method further includes bonding the replacement core with theairfoil, wherein during bonding, the replacement core and the airfoilare chemically bonded to one another along at least a portion of thejoint and the one or more plies are chemically bonded to one another andat least one of the plies is chemically bonded to both the airfoil andthe replacement core. In such implementations, bonding can includeinserting the joined airfoil and replacement core and the one or moreplies overlaying the joint into an autoclave for a predetermined time oruntil the joined components reach a predetermined temperature. Bondingcan also include inserting the joined airfoil and replacement core andthe one or more plies overlaying the joint into a furnace for apredetermined time or until the joined components reach a predeterminedtemperature. The joined components can be chemically bonded at theirmatrix cured interfaces, for instance.

In some implementations, prior to overlaying the joint with the one ormore plies, the method further includes forming one or more coolingholes in the replacement core. For example, the replacement core 130 canhave cooling holes 148 preformed or already defined in replacement core130 prior to being overlaid by plies 170. This can, for example, reducescrap cores and plies as the holes can more easily be formed beforesubjecting the repaired airfoil to one or more bonding processes.

In some implementations, the airfoil is formed from a CMC material andthe airfoil is configured for use in a gas turbine engine. In someimplementations, the airfoil is formed from a PMC material and theairfoil is configured for use in a gas turbine engine.

FIG. 12 provides a flow diagram of an exemplary method according to anexemplary embodiment of the present subject matter. In particular, FIG.12 provides a flow diagram for a method for repairing a component formedfrom a composite material, such as e.g., a CMC or PMC material.

At (402), the method (400) includes joining a replacement material withthe component, wherein the component comprises an interlocking featureand the replacement material comprises an interlocking featurecomplementary to the interlocking feature of the component, wherein whenthe replacement material is joined with the component, the interlockingfeatures of the replacement material and the component interlock to forma joint. For instance, the component can be the airfoil 100 and thereplacement material can be the replacement core 130 illustrated anddescribed herein. The interlocking feature for the component can be theinterlocking feature 116 shown and described herein for airfoil 100. Forexample, the interlocking feature 116 can be dovetails 118 (rounded asshown in FIG. 4 or flaring as shown in FIG. 8), finger projections 150as shown in FIG. 9, fir trees 154 as shown in FIG. 10, a combination ofthe foregoing, etc. Likewise the interlocking feature of the replacementmaterial can be the interlocking feature 144 shown and described hereinfor replacement core 130.

At (404), the method (400) includes overlaying the joint with one ormore plies. For instance, one or more plies 170 as shown in FIGS. 6 and7 can overlay the joint 160 formed by the interlocking features 116, 144of the airfoil 100 and replacement core 130, respectively. By overlayingplies 170 over the joint 160, the airfoil can be rebuilt or recontouredto a predetermined geometry or desired shape, the plies 170 can provideimproved shear and tensile strength to the repaired airfoil and act toseal the joint 160.

At (406), the method (400) includes bonding the component with thereplacement material and the one or more plies with the component andthe replacement material. For example, after overlaying the joint withone or more plies, the method further includes bonding the replacementcore with the airfoil, wherein during bonding, the replacement core andthe airfoil are chemically bonded to one another along at least aportion of the joint and the one or more plies are chemically bonded toone another and at least one of the plies is chemically bonded to boththe airfoil and the replacement core. In such implementations, bondingcan include inserting the joined airfoil and replacement core and theone or more plies overlaying the joint into an autoclave for apredetermined time or until the joined components reach a predeterminedtemperature. Bonding can also include inserting the joined airfoil andreplacement core and the one or more plies overlaying the joint into afurnace for a predetermined time or until the joined components reach apredetermined temperature.

In some implementations, the component is formed from a CMC material andis configured for use in a gas turbine engine. In some otherimplementations, the component is formed from a PMC material and isconfigured for use in a gas turbine engine. In addition, exemplaryimplementations described above with reference to method (300) areequally applicable to method (400).

This written description uses examples to disclose the invention,including the best mode, and also to enable any person skilled in theart to practice the invention, including making and using any devices orsystems and performing any incorporated methods. The patentable scope ofthe invention is defined by the claims and may include other examplesthat occur to those skilled in the art. Such other examples are intendedto be within the scope of the claims if they include structural elementsthat do not differ from the literal language of the claims or if theyinclude equivalent structural elements with insubstantial differencesfrom the literal language of the claims.

What is claimed is:
 1. A method for repairing a component having adamaged region, the component formed from a composite material anddefining a span length that extends from a first end of the component toa second end of the component, and a reference line that is definedmidway between the first end and the second end, the damaged regionbeing closer to the second end than the first end, wherein the componentis an airfoil for use in a gas turbine engine, the method comprising:joining a replacement material with the component, the replacementmaterial having a tip end, a connection end, a leading edge, and atrailing edge, wherein the component comprises an interlocking featureand the connection end of the replacement material comprises aninterlocking feature complementary to the interlocking feature of thecomponent, wherein when the replacement material is joined with thecomponent, the interlocking features of the replacement material and thecomponent interlock to form a joint that defines a joint width and ajoint length, the joint width being perpendicular to the joint lengthand the joint length spanning a width of the component; overlaying thejoint with a plurality of plies, wherein overlaying the joint compriseslaying a first ply over the joint on a pressure side of the airfoil andlaying a second ply over the joint on a suction side of the airfoil; andbonding the component with the replacement material and bonding theplurality of plies with the component and the replacement material;wherein the joint is only between the reference line and the first endof the component; wherein the first ply extends beyond the leading edgeof the replacement material; wherein overlaying the joint with aplurality of plies comprises laying one or more ply layers on thepressure side of the airfoil and laying one or more ply layers on thesuction side of the airfoil, wherein the number of ply layers on thepressure side of the airfoil does not equal the number of ply layers onthe suction side of the airfoil.
 2. The method of claim 1, wherein atleast one of the plurality of plies extends from a center of the jointwidth in a first direction a distance equal to the joint width andextends from the center of the joint width in a second direction adistance equal to the joint width, the first direction opposite thesecond direction.
 3. The method of claim 1, wherein the component isformed from a polymer matrix composite (PMC) material.
 4. The method ofclaim 1, wherein the interlocking feature of the component is one ormore dovetails and the interlocking feature of the replacement materialis one or more dovetails.
 5. The method of claim 1, wherein theinterlocking feature of the component and the interlocking feature ofthe replacement material are configured to counteract tensile forces. 6.The method of claim 1, wherein the interlocking feature of the componentand the interlocking feature of the replacement material are configuredto counteract shear forces.
 7. The method of claim 1, wherein the stepof bonding comprises inserting the component, the replacement material,and the plurality of plies in an autoclave to cause the component tochemically bond with the replacement material and the plurality of pliesto chemically bond with the component and the replacement material. 8.The method of claim 1, wherein the first ply mates with the second ply.9. The method of claim 1, wherein at least one of the plurality of plieshas a different length than another one of the plurality of plies. 10.The method of claim 1, further comprising machining the component andthe replacement material after boding the component with the replacementmaterial.